(1) Field of the Invention
The invention relates to gas turbine engines, and more specifically to a cooled fluid sealing arrangement disposed between blades and vanes of such engines.
(2) Description of the Related Art
Gas turbine engines operate by compressing ambient air with a forward compressor, injecting a fuel, burning the air-fuel mixture in a central combustor and converting the energy of combustion into a propulsive force. Combustion gases exit the combustor through an annular duct, where the gases drive one or more axial stages of circumferentially distributed turbine blades. Each bladed stage transfers the combustion gas energy to a rotor attached to a central, longitudinal shaft. Interposed with the rotating blade stages are stationary vane stages affixed to radially outer casing structures, circumscribing the rotor. Two or more rotors may operate independently of one another and at differing speeds via concentric shafts. Gas turbine engines are flexible power plants that are typically used for powering aircraft, ships and generators.
In order to withstand combustion gas temperatures that regularly exceed 2000 degrees Fahrenheit and pressures exceeding 400 pounds per square inch absolute, turbine components such as blades, vanes and seals are cooled with lower-temperature, higher-pressure cooling air. The cooling air is bled from the compressors, then directed axially rearward and radially inward of the rotors to the turbine components, bypassing the combustor altogether. Once delivered to the turbine, a significant portion of the cooling air is directed radially outward to the blades, vanes and seals by the centrifugal force of the turning rotors. In order to achieve the greatest heat reduction benefit from the cooling air, the interfaces of the rotating blade stages and stationary vane stages must be effectively sealed.
The interfaces of the rotating blade stages and stationary vane stages are particularly difficult to seal due to the differences in thermal and centrifugal growth between the rotors and the cases. The high relative speeds, extremely high temperatures and pressures also present seal design challenges in the turbines. In the past, designers have attempted to seal the interfaces of the rotating blade stages and stationary vane stages with varying degrees of success.
An example of such a turbine seal is a labyrinth seal. In a typical blade to vane interface, a multi-step labyrinth seal, comprising stationary lands and rotating runners or knife-edges, restricts leakage of the cooling air radially outward, into the combustion gases. The runners project from annular supports, which are typically fastened to the rotor with bolted flanges and/or with snap fits. The supports are independent components, adding to the manufacturing costs and complexity of the turbine. The supports also contribute additional rotational mass to the rotors, which reduces the engine-operating efficiency. Also, the attachments at the interfaces of the supports and the rotors create an additional leakage path for the cooling air. Placement of the supports is influenced by adjacent components and typically does not optimize the distribution of the cooling air.
What is needed is a blade to vane interface seal that doesn't require separate seal support components, and also improves the apportioning of cooling air to the seal itself.